In recent years, as a liquid-fueled engine which is mounted in an aerospace vehicle, such as a rocket, the use of a turbo pump-type rocket engine in which a propellant (for example, liquid hydrogen as fuel and liquid oxygen as an oxidizer) is pumped to a combustor by a turbo pump to obtain a large propulsive force has become mainstream. For example, Patent Document 1 describes an expander cycle engine, as one form of a turbo pump-type rocket engine, in which liquid hydrogen pumped from a fuel turbo pump is used for regenerative cooling of a combustor and gasified, hydrogen gas is used for driving the fuel turbo pump and the oxidizer turbo pump and then introduced to the combustor, and liquid oxygen is pumped directly to the combustor from the oxidizer turbo pump.
This turbo pump-type rocket engine is attracting attention as a rocket engine for a vertical takeoff and landing aircraft. The vertical takeoff and landing aircraft is designed on the assumption that the aircraft flies using the profile shown in FIG. 8. The profile shown in FIG. 8A has Ph1: vertical lift, Ph2: pitch maneuver, Ph3: MECO (Main Engine Cut-Off), Ph4: wide range (hovering), Ph5: re-entry/lift flight, Ph6: approach guide, Ph7: engine resignation, Ph8: landing guide, and Ph9: vertical landing. In Ph4, in addition to wide range (hovering), for example, ballistic flight or orbital flight may be carried out. For this reason, unlike a typical disposable rocket, in a rocket engine which is mounted in a vertical takeoff and landing aircraft, high-speed responsiveness (response frequency equal to or higher than 1 Hz) and wide range thrust variability are required during operation of the rocket engine from controllability against crosswinds during landing or thrust throttling corresponding to the body weight which becomes half that of launching during landing. Above all, high-speed responsiveness and wide range thrust variability are required in the range of Ph1 to Ph2 and Ph7 to Ph9 of FIG. 8A, in particular, in the range of Ph7 to Ph9 during landing.
A conventional rocket engine is designed on the assumption of a one-way operation to the space. In general, in order to minimize a gravity loss, the rocket engine is operated with the maximum thrust at the time of launching, and then the thrust is simply quasi-statically squeezed little by little from the restrictions of body acceleration, aerodynamic load, and the like. That is, the conventional rocket engine is designed on the assumption that the characteristics in the substantially normal state are evaluated, and in general, does not take into consideration thrust responsiveness. This is also applied to a turbo pump-type rocket engine.
FIG. 8B shows the combustion test result of a conventional turbo pump-type rocket engine. In FIG. 8B, the horizontal axis represents time (sec), and the vertical axis represents combustion pressure Pc(kg/cm2), the number of rotations Nf (rpm) of the fuel turbo pump, and the number of rotations No (rpm) of the oxidizer turbo pump. As shown in FIG. 8B, it is understood that it takes time, about five seconds, until the combustion pressure Pc corresponding to the engine thrust is lowered from about 30 (kg/cm2) to about 20 (kg/cm2), that is, until the thrust is changed to 66%. When being converted to a response frequency, the response time of five seconds is 0.2 (Hz), and it is not possible to satisfy high-speed responsiveness which is required in a rocket engine for a vertical takeoff and landing aircraft, making it difficult for fine maneuver during landing. From FIG. 8B, it is understood that the numbers of rotations Nf and No of the respective turbo pumps are also changed with a change in the thrust (a change in combustion pressure Pc).